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. Aspect Ratio ,AR
I.:. A. C., t{ft.)
Root Chord ( ft. )
GALCIT
Report //119 Page 6.
II. IJ!ethod of i1Iakinfl' Tests and Calculations and of Presenting Results; Notation.
The normal experimental setup is indicated schernatically in
Fig. 3. Unfortunately, at the time the tests were made, only five balances
were available. He~e for lift 9 drag, and pitching moment investigations
the two yaw balances were not usedo For rolling and yawing moment tests the
drag balance and one lift ba1.a.11ce were moved into the positions of the two
yaw balances. For this reason lift and drag could not be measured simul.
taneously with rolling and yawing moments. (This point is further discussed
in III, 5 below}.
The tare drag of the pyramid wire systems running to the wing
trunnions (cf. Photo 2) was known r:om previous GALCIT investigations. The
tare drag and moment of the sting, tail wire, and counte2·weight wire were
..... determined by a~ testing a wing alone with an auxiliary sting attached. and
b) completely enclosing the tail and counterweight wires in streamlined
windshields. T'ne tare drag at the high speed attitude of the ail9Plane was
about 65% of the total J;nrasite drag at this attitude. The tare moment was
extremely small at all angles cf attack.
All drags, angles of attack, and pitching moments were corrected
by the Prandtl theory of -~unnel wall interference to give fre~ air condi·tiona.
Rolling and yawing moments and side force were uncorrected for the effect
of wall interferenceo
All observations were reduced to the standard American system of
absolute units (the notation for rolling moment, yawing moment, and side force
is different from that rec arrunended by the N.A.C.A.)
Lift CD - Dra
CM - Stall in Moment
Cr, -- .. ... f v2s ev2s ev2s t
2 2
Side Force ... Rollin l\1oment
Cy - Yawi Moment c - ... Cr -
s ev2s ev2s b ev2s b
2 2 2
GALCIT
Report ;,tug
alone observations.
Page 8.
In ma%ing performance estimates and in comparing the effects
of various modifi cations the equivalent parasite area has been used, where
f :s equ.i vale.at p:lrasi te area -- a.rasite dr,,
ev2
2
s Cllp S
Certain additional symbols are used in Section III which are
there defined.
It should be mentioned that the plotted e:A.1>erimental points
represent di1•ect observations with no fairing, except that the tare drag
results were faired before being subtracted :from the observed total drags
to give ·the final values.
III. EJ...rnorimental Results and Discussion.
In view of the tremendous n~er of individual observations all
results are given in the form of plotted exwrimental points and faired
curves, practically no tabular data being given in the Report. Such detailed
t abular data are nvail.able in the files or the GALCIT. The results are
discussed below in accordance with the grouping previously mentioned.
l) Three Component Measurements on Wi s Alone Fi :.os. 4 5)
The results for the four wings reduced to infinite aspect ratio
conditions are plotted in .l!'ig. 4. CLm~ for vr,ing WJ. is 4 or 5 percent lc;1er
than that for the other three which attai11 the quite normal values (for the
GALCIT tunnel) of between 1.30 and 1.32. The rather low value of Cv
0
fo:r
wing Wt is not especially surprisir..g when it is considered that Wt was ob-tained
from W0 by adding on large tips of comparatively thin section. The
rather cUl.. ious shape of the Cn0 curve between Ct C4 0.3 and L 0 for i;;ing W2
seemed a little suspicious when the results were finally plottedQ Hence it
was decided to make the series of Runs including 192 as a check on t he
earlier series including 152~ Run 192 was tal~ en ten weeks u:fter Run 152,
duriri.g which time the bala..'TlC e sys ter.i had been cor.ipletely shifted twice and
the model completely disassembled and reassa~bled. The aereement of t he t wo
GA:CIT
Re~rt }ll9 Pa.::;e 1·3.
In 7ie~7 of t:1c: uns::i.tisfactory stability ch~s:-a.cte1i.ztics of ..:;1:7t ,
a second will[; ~72 , o.nd a new set of horizontal t ail surfaces 112, \7crc coil-struc
ted , as has already been men tionedo Unfortunately t he stabilizer of
H2 was built rigidly into the fuselage at a setting of o0
, so that e:i..-peri-ments
with wi ng ':{2 and horizontal tail surfaces removed could not be mo.deQ
Al so, because of the elaborate set of tests with S1Wt 1 no compl ete ir~ves ti ga - ~
ti on of elevator and fl ett ner effect iveness was undertaken vri t h S2B'2 •
In add.i t i on, the elevators of H2 were not stntically balanced so that no
elevator free measurements could be ma de~ 'i.'he pitching mor-1ents for wing W2
alone and for s2w2 ( i. e o stabilizer o0 , elevator o0 , f lettner o0 } are
plotted in ll'ig~ 9o It is seen that the stabi lity of S2 ~72 is -;er·J t;ood over
. dCMt •
the norn.al fl y:i.ng range ~ dCL = -0.15} while eve:i at t he s tall t here is no
actual instability ~
The ef'fecti veness of elevutor and flettner in permitting t rim at
the stall is discussed later in connection with the measurer.lents on bot t on
surface flnps \section 6)" CM vs. Ct l1Urvcs are given in Fie . 34 i n this
• connection fo1, s2w2 and for s2w2w2 ii s2w2 • Compari ng the tno cur1res one sees
that ·the addition of' w2 decreases t h e stability very slightly a..J.d m~e s the
plane approximatel y ne1itrally stable for quite a ran~e of a:rlBles of attack
a·t the stall., The curves of Fig. 33 for elevator up indicate ample control
to enable t he tail to be gotten down for landing . The effect of i"'laps on
s tabil ity and trim is discussed in section 6 helo'if.
5) Lateral { i 11aludi ~ Directional) Control
As mentioned in section II, the nurnber of wind tunnel balances
available at the t t--·c of the t ests was i nsuf'fi c:1 ent to per r:li t six components
to be measured sirmltc.neousl:ro ':lith the nornnl "lift, drag, pitchine; moment"
setup the rollin&: moment about an axis through the dillillond shaped suspension
frame (cf . FiG• 3) could be measured, It i1as at first assumed that this was
identical 77i t 11 the roll in::: rii.o;;icnt about the fuselage axis and a series of
GALCIT REPORT PAGE 29 .
TO STATIC RINGS
IN ENTRANCE AND
WORKING SECTIONS
BALANCE ROOM
TO SELSYN GENERATOR
OWG. WT26
WIND
TUNNEL
/
/
,
SIX COMPONENT SETUP FOR TEN FOOT WIND TUNNEL TESTS
AT GUGGENHEIM AERONATICS LABORATORY
CALIFORNIA INSTITUTE OF TECHNOLOGY
FIG. 3 .
/
/,.//
/'
":~~- :-~ ~ ~;:. 1 ".·?
----j .'.~-<1~
l. - ., •
~~: -~ ~~ ~7 l '"· I:·: , f-:-~
=:·: .. ~: ( .
--.- ·-·-1I ' :·. : . 1 i: .
- - --- :r-:-:-
• _. I • - .,....r ..... , • ~
S!.L.: ·=>r )-~ :=-:
.. i -: - -:_" .
. ~ . . - '
: : ~ . i.~~~ -~ ~ ~ - :~
·1 ~ •; • • .·• - • . • ~ I
ti ·1 t . - ::-:
-- - - - -- - -· .
. ~l . -: l~-~ ~~~:
. -- I-. - - .. I, . -
-- ' .-;, L+' :~~ j _;~~-
_-_ -;:~ ~~ i-: <-~-~
a:-:.: -=-- -,- -- -
. ~-.. ..
~J?- --=-_:-)~::_
- - -- - -~ - - - !---:_:~_-:- -
·_ - :.=-:- .. _ _:--__ r--.:~­-
.:..--::-_: - - -- -
"'°-:- _t _-
-- -- ---
•

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. Aspect Ratio ,AR
I.:. A. C., t{ft.)
Root Chord ( ft. )
GALCIT
Report //119 Page 6.
II. IJ!ethod of i1Iakinfl' Tests and Calculations and of Presenting Results; Notation.
The normal experimental setup is indicated schernatically in
Fig. 3. Unfortunately, at the time the tests were made, only five balances
were available. He~e for lift 9 drag, and pitching moment investigations
the two yaw balances were not usedo For rolling and yawing moment tests the
drag balance and one lift ba1.a.11ce were moved into the positions of the two
yaw balances. For this reason lift and drag could not be measured simul.
taneously with rolling and yawing moments. (This point is further discussed
in III, 5 below}.
The tare drag of the pyramid wire systems running to the wing
trunnions (cf. Photo 2) was known r:om previous GALCIT investigations. The
tare drag and moment of the sting, tail wire, and counte2·weight wire were
..... determined by a~ testing a wing alone with an auxiliary sting attached. and
b) completely enclosing the tail and counterweight wires in streamlined
windshields. T'ne tare drag at the high speed attitude of the ail9Plane was
about 65% of the total J;nrasite drag at this attitude. The tare moment was
extremely small at all angles cf attack.
All drags, angles of attack, and pitching moments were corrected
by the Prandtl theory of -~unnel wall interference to give fre~ air condi·tiona.
Rolling and yawing moments and side force were uncorrected for the effect
of wall interferenceo
All observations were reduced to the standard American system of
absolute units (the notation for rolling moment, yawing moment, and side force
is different from that rec arrunended by the N.A.C.A.)
Lift CD - Dra
CM - Stall in Moment
Cr, -- .. ... f v2s ev2s ev2s t
2 2
Side Force ... Rollin l\1oment
Cy - Yawi Moment c - ... Cr -
s ev2s ev2s b ev2s b
2 2 2
GALCIT
Report ;,tug
alone observations.
Page 8.
In ma%ing performance estimates and in comparing the effects
of various modifi cations the equivalent parasite area has been used, where
f :s equ.i vale.at p:lrasi te area -- a.rasite dr,,
ev2
2
s Cllp S
Certain additional symbols are used in Section III which are
there defined.
It should be mentioned that the plotted e:A.1>erimental points
represent di1•ect observations with no fairing, except that the tare drag
results were faired before being subtracted :from the observed total drags
to give ·the final values.
III. EJ...rnorimental Results and Discussion.
In view of the tremendous n~er of individual observations all
results are given in the form of plotted exwrimental points and faired
curves, practically no tabular data being given in the Report. Such detailed
t abular data are nvail.able in the files or the GALCIT. The results are
discussed below in accordance with the grouping previously mentioned.
l) Three Component Measurements on Wi s Alone Fi :.os. 4 5)
The results for the four wings reduced to infinite aspect ratio
conditions are plotted in .l!'ig. 4. CLm~ for vr,ing WJ. is 4 or 5 percent lc;1er
than that for the other three which attai11 the quite normal values (for the
GALCIT tunnel) of between 1.30 and 1.32. The rather low value of Cv
0
fo:r
wing Wt is not especially surprisir..g when it is considered that Wt was ob-tained
from W0 by adding on large tips of comparatively thin section. The
rather cUl.. ious shape of the Cn0 curve between Ct C4 0.3 and L 0 for i;;ing W2
seemed a little suspicious when the results were finally plottedQ Hence it
was decided to make the series of Runs including 192 as a check on t he
earlier series including 152~ Run 192 was tal~ en ten weeks u:fter Run 152,
duriri.g which time the bala..'TlC e sys ter.i had been cor.ipletely shifted twice and
the model completely disassembled and reassa~bled. The aereement of t he t wo
GA:CIT
Re~rt }ll9 Pa.::;e 1·3.
In 7ie~7 of t:1c: uns::i.tisfactory stability ch~s:-a.cte1i.ztics of ..:;1:7t ,
a second will[; ~72 , o.nd a new set of horizontal t ail surfaces 112, \7crc coil-struc
ted , as has already been men tionedo Unfortunately t he stabilizer of
H2 was built rigidly into the fuselage at a setting of o0
, so that e:i..-peri-ments
with wi ng ':{2 and horizontal tail surfaces removed could not be mo.deQ
Al so, because of the elaborate set of tests with S1Wt 1 no compl ete ir~ves ti ga - ~
ti on of elevator and fl ett ner effect iveness was undertaken vri t h S2B'2 •
In add.i t i on, the elevators of H2 were not stntically balanced so that no
elevator free measurements could be ma de~ 'i.'he pitching mor-1ents for wing W2
alone and for s2w2 ( i. e o stabilizer o0 , elevator o0 , f lettner o0 } are
plotted in ll'ig~ 9o It is seen that the stabi lity of S2 ~72 is -;er·J t;ood over
. dCMt •
the norn.al fl y:i.ng range ~ dCL = -0.15} while eve:i at t he s tall t here is no
actual instability ~
The ef'fecti veness of elevutor and flettner in permitting t rim at
the stall is discussed later in connection with the measurer.lents on bot t on
surface flnps \section 6)" CM vs. Ct l1Urvcs are given in Fie . 34 i n this
• connection fo1, s2w2 and for s2w2w2 ii s2w2 • Compari ng the tno cur1res one sees
that ·the addition of' w2 decreases t h e stability very slightly a..J.d m~e s the
plane approximatel y ne1itrally stable for quite a ran~e of a:rlBles of attack
a·t the stall., The curves of Fig. 33 for elevator up indicate ample control
to enable t he tail to be gotten down for landing . The effect of i"'laps on
s tabil ity and trim is discussed in section 6 helo'if.
5) Lateral { i 11aludi ~ Directional) Control
As mentioned in section II, the nurnber of wind tunnel balances
available at the t t--·c of the t ests was i nsuf'fi c:1 ent to per r:li t six components
to be measured sirmltc.neousl:ro ':lith the nornnl "lift, drag, pitchine; moment"
setup the rollin&: moment about an axis through the dillillond shaped suspension
frame (cf . FiG• 3) could be measured, It i1as at first assumed that this was
identical 77i t 11 the roll in::: rii.o;;icnt about the fuselage axis and a series of
GALCIT REPORT PAGE 29 .
TO STATIC RINGS
IN ENTRANCE AND
WORKING SECTIONS
BALANCE ROOM
TO SELSYN GENERATOR
OWG. WT26
WIND
TUNNEL
/
/
,
SIX COMPONENT SETUP FOR TEN FOOT WIND TUNNEL TESTS
AT GUGGENHEIM AERONATICS LABORATORY
CALIFORNIA INSTITUTE OF TECHNOLOGY
FIG. 3 .
/
/,.//
/'
":~~- :-~ ~ ~;:. 1 ".·?
----j .'.~-<1~
l. - ., •
~~: -~ ~~ ~7 l '"· I:·: , f-:-~
=:·: .. ~: ( .
--.- ·-·-1I ' :·. : . 1 i: .
- - --- :r-:-:-
• _. I • - .,....r ..... , • ~
S!.L.: ·=>r )-~ :=-:
.. i -: - -:_" .
. ~ . . - '
: : ~ . i.~~~ -~ ~ ~ - :~
·1 ~ •; • • .·• - • . • ~ I
ti ·1 t . - ::-:
-- - - - -- - -· .
. ~l . -: l~-~ ~~~:
. -- I-. - - .. I, . -
-- ' .-;, L+' :~~ j _;~~-
_-_ -;:~ ~~ i-: